Stability and control augmentation system

ABSTRACT

A stability and control augmentation system for a craft wherein a stabilizing signal is generated in response to external disturbances to activate a high gain stability feedback loop to introduce controlled angular rate. An operator control input activates both the stability feedback loop circuit and a control feedforward loop circuit to change the reference point for the feedback loop without altering the gain thereof. The high gain of the feedback loop dominates the response characteristics of the craft and provides effectively constant response characteristics through the operational envelope. The control feedforward loop circuit alters the reference point for feedback loop circuit so that the mechanical control input may be introduced to provide the desired control response.

I United States Patent [1113584314 [72] Inventor Milford R. Murphy3,082,979 3/1963 Hendrick 244/77 (0V) Arlington, Tex. 3,237,107 2/1966Bresenoffet al 244/77 (D) [21 P 12 Primary Examiner-Milton Buchler 5'' dJuly 1971 Assistant Examiner-Jeffrey L. Forman [4 1 {:31 CAttorneys-Richards, Harris & Hubbard, v. Bryan Medlock, [73 1 "08pm Jr., Harold E, Meier and George Galerstein Hurst, Tex.

ABSTRACT: A stability and control augmentation system for [54] STABILITYAND CONTROL AUGMENTATION a craft wherein a stabilizing signal isgenerated inresponse to SYSTEM external disturbances to activate a highgain stability feedback 6 claims 10 D loop to introduce controlledangular rate. An operator control "wing .88- input activates both thestability feedback loop circuit and a [52] US. Cl. 244/77 control f df dloop circuit to change the reference point [51] Cl i B649 13/18 for thefeedback loop without altering the gain thereof. The [50] Field ofSearch ..3 l 8/20.245; high gain f the f db k loop dominates theresponse charac 244, 77 77 teristics of the craft and provideseffectively constant response characteristics through the operationalenvelope. The control [56] References cued feedforward loop circuitalters the reference point for feed- UNITED STATES PATENTS back lo'opcircuit so that the mechanical control input may be 2,998,210 8/1961Carter 244/77 (D) introduced to provide the desired control response.

, Y fIG 44 48 CONTROL P i 32 i so l f f f TRANSDUCER ,fgggl i g R ATEGYRO a TRANSDUCER 38 28 54 24 COMPENSATING A P [26 COM PENSATWG NETWORKl i NETWORK ATENTED JUN] 5 l97l SHEET 1 OF 4 O O 4 8 W 2 R 2 K 4. A f Y[AR R 6 8% F ii- E WT R T m A M R w C q 4 I L fl w e W 2 G T S W I C8 F5 R MQ [.OCC 4 T U 5 ABD I T T I DS T m H AFR u T l 4 W m N w A T A U UA n 2 a G [MR r D mm S ET 0 M PE m W C O FIG. 4

DISTURBANCE FIG. 3

INVENTOR MILFORD R. MURPHY ATTORNE Y FIG. 6

FREQUENCY IN RADIANS PER SECOND GAIN IN db mamimumsmn 35841814 SHEET 3UF 4 I FIG. 8

ovv\,- l N\r- E FIG. 9

FREQUENCY lN RADIANS PER SECOND STABILITY AND CONTROL AUGMENTA'I'IONSYSTEM THE INVENTION This invention relates to stability augmentationsystems and more particularly to systems in which there is augmentationof both stability and of control to improve the normal, inherentproperties of the controlledbody.

THE PRIOR'ART Existing systems directed at the augmentation of stabilitycharacteristics of an aircraft, as opposed to systems directed atestablishment and maintenance of a flight attitude, possess severalsignificantly objectional qualities. For example, aircraft systemsgenerally have employed a feedback signal which opposes-the pilotcontrol input as well as the external disturbing functions (i.e., windgust, weapon recoil, etc.). To avoid problems involved in suchoperation, the gain of the feedback loop of such system is set at alower or compromised value. This has. compromised both the stability andcontrol characteristics of the aircraft.

Some such relatively simple systems operate in compromise of the desiredstability control characteristics by employing an effectively constantfeedback gain throughout the aircraft flight envelope. Such operation isundesirable in some instances. For example, the optimum feedback gainfor hovering a helicopter would be too high for high speed forwardflight.

In more complex systems, alteration of the gain to accommodate changingflight conditions is achieved by intervention of the pilot who canmanipulate a control to change the feedback gain (and/or phasing). Instill other systems, the feedback gain and phasing is alteredautomatically in response to sensed changes in the flight condition,principally changes in speed. These systems contribute to theimprovement of the overall stability and control operatingcharacteristics. However, they possess a basic deficiency in that theycompromise control and stability.

Relatively simple systems employ a fixed compromise between control andstability. More complex systems, in response to changing flightconditions, employ an automatically varying optimum compromise.

A system in accordance with this invention represents a significantadvance over prior systems in that no compromise between stability andcontrol is required. Stability characteristics and control responsecharacteristics are optimized through the entire significantcontrol/response range of the aircraft.

The significant control/response range of a pilot/airframe system isfixed at its lower end by the reaction time of the pilot at which he isfirst capable of intelligently analyzing aircraft movement. At the upperend, the range is fixed by the time at which the pilot must begin torely upon memory in order to analyze the aircraft motion.

SUMMARY throughout the significant control/response range.

THE DRAWINGS A more complete understanding of the invention and itsadvantages will be apparent from the specification and claims and fromthe accompanying drawings illustrative of the invention.

FIG. 1 is a block diagram of a stability and control augmentation systemin accordance with the present invention;

FIG. 2 is a block diagram of a system functionally equivalent to thesystem of FIG. I with the electrical portion of the pilot inputdisconnected;

FIG. 3 is a block diagram of a system that is functionally equivalent tothe system of FIG. 1 with the pilots mechanical input disconnected;

FIG. 4 is a diagrammatic representation of the parallel relationship ofthe systems of FIGS. 2 and 3;

FIG. 5 is a Bode plot illustrating the relationship between.

the transfer functions of an airframe and its stabilizing system;

' FIG. 6 is a Bode plot illustrating the stability transfer function ofa helicopter airframe augmented by the system of this invention;

FIG. 7 is a schematic diagram of a stability feedback loop compensatingnetwork;

FIG. 8 is a schematic diagram of a feedforward loop compensatingnetwork;

FIG. 9 is a Bode plot illustrating the control transfer function of ahelicopter airframe augmented .by a system in accordance with thisinvention; and

FIG. 10 is a schematic of a preferred embodiment of this invention.

THE PREFERRED EMBODIMENTS In FIG. 1, a helicopter control andstabilization system is shown in configuration to embody the presentinvention and includes an inner control loop 4, a stability feedbackloop 6, and a feedforward loop 8.

Included in the above system is an actuator 16 of a limited authority,series type which is electrohydraulically actuated. An actuator feedbacktransducer 50 is an electromechanical position feedback transducer forthe electrohydraulic actuator 16. A rate gyro 20 senses attitude rateand provides the basic feedback signal. Components 16, 20 and 50 arewell-known standard control hardware found in prior helicopter controlsystems.

A stability compensating network 24 employed herein is an electricalnetwork complementing the rate gyro 20 within the feedback loop 6 toprovide desired stability characteristics of the airframe 48 throughoutthe flight envelope.

A control compensating network 38 is an electrical network within thepilot feedforward loop 8 which operates on the electrical signalgenerated by transducer 32 upon movement of a control stick 10.

Control system 44 consists of standard control means for commanding achange of direction or attitude of flight of the airframe 48. It may beany one of the main rotor lateral control system, main rotorlongitudinal control system or the tail rotor control system of ahelicopter acting to control roll, pitch and yaw respectively. Thus forthe present description'a single axis stability augmentation controlsystem will be described, it being understood that such a system willgenerally be employed for each of three axes when applied to aircraftsuch as a helicopter.

The system comprises two basic loops, the stability feedback (outer)loop 6 and the pilot feedforward loop 8, with a commmon inner controlloop 4. To facilitate understanding of the entire system, the two basicloops and the inner control loop will first be considered relativelyindependently.

INNER CONTROL LOOP The control or inner loop 4 comprises an amplifier 28and an actuator 16 mechanically connected by means 42 to the actuatorfeedback transducer 50. An input electrical signal is applied toamplifier 28, as for example, through electrical connection 26 of thestability loop 6, and is amplified and applied to a torque motor windingin the hydraulic actuator 16 thereby causing the actuator to extend orretract, depending upon the polarity of the input signal to theamplifier. The extension or retraction of actuator 16 is detected by thetransducer 50 which provides an opposing electrical signal to amplifier28 through connection 54. Actuator 16 will continue to position thecontrol system 44 until the input signals to amplitier 28, throughconnections 26 and 54, are equal and cancel each other. The magnitude ofthe control achieved by means of the control system 44 per unit of inputto amplifier 28 from the channel 26 is the static closed loop gain ofthe control loop 4 and its value is determined the design of theairframe loop as will be hereinafter shown.

STABILITY FEEDBACK LOOP The stability feedback loop 6 comprises theairframe 48, rate gyro 20, stability compensating network 24, the innercontrol loop 4 (28-16-50) and the control system 44. The basic functionof the stability feedback loop 6 is to provide for a positioning of thecontrol system 44 that opposes external disturbances of airframeattitude. Except for the characteristics of the stability compensatingnetwork 24, loop 6, as an entity, is conventional. Network 24 comprisesa washout network and a lag lead-type network which shall be discussedhereinafter.

In operation, a movement of the airframe 48 causes the rate gyro 20 togenerate on electrical signal which is transmitted through compensatingnetwork 24 and the amplifier 28 to the actuator 16. Actuator 16 willthus be energized to extend or retract. As described in connection withthe inner control loop, actuator 16 actuates the control system 44 toclose the loop.

PILOT F EEDF ORWARD LOOP The pilot feedforward loop 8 comprises a pilotcontrol member 10, control motion transducer 32, control compensatingnetwork 38, and inner loop 4. Movement of the member 10 produces aproportionate change in electrical output from the mechanicallyconnected transducer 32 which output is acted on by network 38. Thenetwork 38 produces a signal that combines with the pilot's mechanicalinput to actuator 16 to effect a change in the control system 44. Theeffect of compensating network 38 is to make the feedback loop gainfrequency sensitive.

ANALYSIS Feedback loops are used primarily to stabilize the output of asystem to a reference or to change the characteristics of a system toobtain a specific response to an input. These two functions of the loopare not mutually exclusive and stabilizing characteristics will affectthe nature of the control response, and vice versa. As above noted, ifstability is considered to be of prime importance, the loop is normallyclosed with a comparatively high feedback gain so that the dynamic andstatic error of the output relative to a reference (usually the input)will be minimized. If control response is considered to be of primeimportance, the loop is preferably closed with a comparatively lowfeedback gain. Where, as in most cases, both stability and control areconsidered substantially equally important, the loop is usually closedwith a gain of intermediate or compromise value.

In accordance with the invention, stability and control may beapproached and controlled separately. More particularly, referring toFIG. 2, the block diagram illustrates a typical closed loop controlsystem, where 0, is an input signal, ill is the system output, G is thetransfer function of the basic airframe, H is the transfer function ofthe feedback loop, and G is the closed loop transfer function which maybe described mathematically as follows:

If G,H is less than I, then G is approximately equal to G Considering Gas the representative transfer function of an unaugmented airframe, thenif G,H is less than 1, the resultant transfer functions 6;, isapproximately equal to the transfer function of the airframe system,i.e., the response characteristics will be substantially determined thebasic airframe; if G,H is greater than 1, the response characteristicsof the system will be approximately equal to the inverse of the transferfunction of the feedback loop, i.e., the response characteristics willbe substantially determined by the stability feedback loop (H). When G,His much greater than unity, the stability feedback loop characteristicsdominate the airframe.

The manner in which the present invention achieves such a system willnow be qualitatively illustrated with reference first to FIG. 3 where G,is the transfer function of the vehicle (referred to as the unaugme ntedairframe), G, is the transfer function of the pilot's electricalfeedforward loop, 0, is the control input, 0;, is the resultant transferfunction of the closed loop, and lb is the system output. The closedloop is provided with a relatively high feedback gain that dominates theresponse characteristics of the basic airframe and results in anaugmented airframe with relatively standard or constant responsecharacteristics throughout the significant control/response range. G, isthe provided so that the resultant equivalent of the system of FIG. .1,with the electrical portion of the pilot feedforward loop 8 disconnectedleaving operative only the pilots mechanical input by means of member 10to actuator 16 and the stability feedback loop 6 which obviously is theloop that operates on a disturbance input.

TI-Ie system represented in the block diagram of FIG. 3 is thefunctional equivalent of the system of FIG. 1 with the pilots controlmember 10 disconnected from actuator 16, there remaining operative theelectrical portion of the pilot feedforward loop 8 and the stabilityfeedback loop 6. This is the system upon which the pilot electricalcontrol input acts through transducer 32 and network 38.

The systems of FIG. 2 and FIG. 3 include the stability feedback loop 6.This is the loop or circuitry upon which external disturbances act andit is the loop which provides optimum stability characteristicsthroughout the aircrafts significant control/response range.

The systems of FIGS. 2 and 3 can be considered as two parallel systemsupon which the pilot control acts. They can be added to produce aresultant transfer function that represents the relationship of theoutput of the system to pilot control input, or, in other words, thecontrol response of the craft.

Referring to FIG. 4, there is shown a block diagram of the systems ofFIGS. 2 and 3 combined into a single system. G is the transfer functionof the electrical input control system and G is the transfer function ofthe mechanical input control system. Using the approach employed todetermine 6;, of FIG. 2, the input signal, 0,, is related to the outputsignal, III, by the following equation:

feedforward loop and the stability feedback loop). Since G;,, thetransfer function of the stability feedback loop, is determined basedupon consideration for stability purposes, t he transfer function 6,.(electrical portion of the pilot feedforward loop) may furnish thedesired values of the term G,./G,,,. IN OTHER WORDS, G, is a designvariable that is used to control the characteristics of the resultanttransfer function G -l-G, to produce the desired control response.

The procedure followed in achieving the present stabilization andcontrol augmentation system will now be explained and illustrated withreference to yaw stability and control in hovering flight and at knotsforward speed. Well-known feedback control technology will be employedalong with applicable mathematical procedures.

The first step, directed at stability alone, will deal with bothhovering and 100 knot forward speed; the second step, directed atdevelopment of yaw control response, will be limited to the 100 knotforward speed condition.

STEP NO. I

The primary objective is to provide a stability feedback transferfunction H (see FIG. 2) that will provide a well damped airframethroughout the flight envelope. Referring to FIG. 5 there is shownsimplified transfer functions of the response of a basic airframe toexternal disturbances for hovering and 100 knot forward speed andplotted on conventional Bode plot (less phase) by lines G and Grespectively. Bode plots involve loop gain (in db.) as a function offrequency plotted in radions per second.

Where the transfer functions G,H is less than one, the resultant closedloop transfer function is approximately equal to G,, i.e., dominated bythe unaugmented airframe. Where the transfer function G,I-I is greaterthan one, the resultant closed loop transfer function is approximatelyequal to lI-I, i.e., dominated by the feedback loop or the augmentedairframe. Thus, it will be appreciated that astability feedback loophaving a gain greater than one through its entire flight spectrum willprovide a system in which the basic stability characteristics of theaircraft will be dominated by characteristics of the feedback loop.

Referring to FIG. 5, the line l/H represents the reciprocal of afeedback transfer function l-I. It will be understood thatrepresentation in terms of l/I-I means that for any given frequency ifthe point on the UH line is lower than the G and G lines, then for thatfrequency, the transfer function G,I-I will be greater than unity andthe response characteristics will be dominated by the characteristics offeedback loop as opposed to the unaugmented airframe. If a point on theUH line is higher than the corresponding points on lines 6,, and G thenthe transfer function 6 H will be less than unity and the responsecharacteristics to external disturbances will be dominated by theunaugmented airframe.

In accordance with this invention, the feedback loop is provided, asshown in FIG. 5, such that the UH line is lower than lines G and G, withvalues that provide the desired stability response throughout thesignificant portion of the flight spectrum.

In practice, the straight line relationships illustrated in FIG. 5 arenot encountered. More representative of actual practice are therelationships illustrated in FIG. 6, which is a Bode plot (less phase)for yaw transfer functions in hovering and 100 knots forward flight fora system embodying the present invention.

Considering line l/I-I, section d-e represents desirable gain values fora given range of frequencies. If section d-e, were extended as astraight line function into the higher frequency region, the stabilityfeedback loop would also dominate at the higher frequencies. This wouldcause the stabilization system to be susceptible to and react to alltypes of high frequency disturbances, including noise, vibrations, etc.Consequently, section g-k, instead of being a straight extension ofsection d-e, is elevated to intersect the G line at it. As a result, theaugmented airframe dominates for 100 knot forward speeds at frequenciesup to approximately 7+ rad/sec. Beyond 7+ rad/sec. the unaugmentedairframe dominates. The point of phasing and require adequate phasemargins for purposes of stability. In a practical application of thisinvention, a phase margin of approximately 60 was maintained from pointsf to Section d-e is not extended as a straight line function into thelower frequencies but is interrupted at point d, as shown. This permitsa more practical circuitry, that is it eliminates the necessity ofexcessively large capacitors, etc. The l/I-I line is maintained beneatngand G lines so that the augmented airframe dominates throughout theentire lower frequency range. A crossover point could well have beenestablished at a frequency below that at which the pilot must begin torelyupon memory for a control action, i.e., around 0.5 rad/sec., and theaugmented airframe would still be dominant throughout the significantcontrol range.

The feedback transfer function H is mechanized by means of the rate gyro20 and network 24 is in the stability feedback loop 6 (FIG. 1). Network24 is made up of a lag network and a washout network. The lag networkoperates on the attitude rate signal to produce the effect from e to k,and the washout network operates on the attitude rate signal to producethe effect from a to e (FIG. 6). The straight line approximation from dto e is representative of the basic gyro attitude rate gain which in oneembodiment was 0.624 of 0, per degree per second of attitude rate.

Referring to FIG. 7, there is shown a representative circuit for network24. Included is a lag network 60 and a washout network 62. The frequencybreaks at e and 3 (FIG. 6) are determined by (R,+R C and R Crespectively, of the lag network; the frequency break at d is determinedby R C of the washout network.

At the upper portions of the frequency spectrum from 1 to 10 radians persec., the lag network 60 operates on the rate gyro signal to maintain arelatively constant phase of 65L-10 over a decade frequency change. Thispermits accommodation of large changes in the transfer function of theairframei.e., changes in airspeed, altitude, gross weight, etc. Anothereffect contributed by use of the lag network is that abrupt externaldisturbances are prevented from introducing equally abrupt correctivecontrol action and there is, instead, a relatively gradual introductionof corrective control tending to reduce pilot discomfort and to providethe aircraft with a more acceptable platform for such functions asgunnery, etc.

The washout network 66 acts to provide a new base reference from whichcontrol changes can be measured, which is of benefit in the lowerfrequencies at which the washout network is activated, because of therelatively long term control inputs involved For example, in hovering,the control pedals are moved by the pilot to an off center position andare retained in that position. Inasmuch as the system is sensitive tochange rather than position," the washout network blocks a steady-staterate signal.

The washout network 66 is the pilot loop and the washout network 62 inthe feedback loop effect cancellation when the pilot introduces acommand and thus the pilot input acts independent of the response of thecontrol loop to disturbance forces. As above stated, it is preferredthat the two washout networks have identical characteristics in order toassure such cancellation.

Thus, the resultant stability response of a typical airframe is definedby line a-c-d-e-f-j for the hovering condition (G and line c-d-e-f-g-h-ifor the 100 knot forward speed condition (G (FIG. 6). It will be notedthat in what is considered the significant control/response range of thepilot/airframe system (c-d-e-f) the stability characteristics will oeessentially the same for both hovering and l knot forward speed and thepilot loop can be provided on this constant to produce specificresponses.

STEP NO. 2

Line l/l-I in FIG. 6 for values of GH l, i.e., from [to 0, representsthe augmented airframe that is operated on by external disturbances and,according to the present invention, is part of the system upon which thepilots mechanical input acts. This is represented by system G, of FIG.4. The pilots electrical input and the stability augmented airframe areconsidered as a separate system G FIG. 4. Also as previously explained,the two systems 6,, and G, are considered in parallel. The present stepin developing the structure and characteristic of 6,. function provisionof a transfer function for the pilots loop in series so that theresultant system will produce a specific response for pilot inputs.

The feedforward loop includes member 10, transducer 32, compensatingnetwork 38 and the inner loop 4. Motion of the control member produces achange in the electrical output signal from the transducer 32 that isproportional to displacement of the control member. This signal isoperated on by compensating network 38 and may be considered as atransfer function G in series with the resultant transfer function G;,.

The present stability and control augmentation system as applied to theyaw system of a helicopter was intended to obtain a constant yaw rate(velocity) per unit of pilot control input. To obtain this particularresponse, network 38 of the pilot control loop should be identical tonetwork 24 of the stability feedback loop. Of course, to obtain otherspecific time constants, as for filtering two cycle per revolutionsignals emanating from the main rotor, differences would be designedinto the networks. However, for this application, network 38substantially cancels the dynamic effects of network 24 leaving the rategyro feedback term which is the controlled variable.

Referring to FIG. 8, there is shown a circuit for network 38, includinga lag network 64 and a washout network 66. This circuit is substantiallythe same as network 24 shown in FIG. 7 except that the shunt resistor inlag network 64 has been omitted to provide a more favorable initialresponse. The resistor R and the capacitor C determine the low frequencybreak points and the resistor R and capacitor C determine the highfrequency break points.

Referring to FIG. 9, the transfer function of the stability augmentedairframe G,,, of FIG. 4 is plotted along with a line (6,.) representingthe desired transfer functions for the pilots electrical inputs. Itshould be appreciated that the situation herein considered is equivalentto that illustrated in the block diagram of FIG. 3 which is thefunctional equivalent of FIG. 1

with the pilot's mechanical input 10 to the actuator 16 discomnected.There remains operative, the electrical portion of the pilot feedforwardloop and the stability feedback loop. Inasmuch as the resultant transferfunction for the pilot's electrical input is the result of themultiplication of the transfer functions of the stability augmentedairframe and of the pilot loop, knowing the first two functions meansthat the transfer function of the pilot loop is readily determinable.

The pilot network is designed to effectively cancel the dynamic effectsof the stability feedback loop network 6 except that the break frequencyof 4 radians per second was deleted so that the network would provideattenuation in the high frequency spectrum and to effect an optimuminitial response.

Having determined G it now remains to sum the transfer functions G and Gwhich summation is approximately indicated (FIG. 9) by the uppermostpoints of the two curves. The resultant transfer function shows that thesystem performs like an integrator for pilot inputs over the frequencyspectrum from 0.8 radians per second to 8.0 radians per second and has atime constant of 0.l25 seconds and an acceleration gain of about 0.64radians per second per second per radian input. So, the mathematicalmodel of the transfer functions for pilot control is:

and the mathematical model of the transfer function on the stabilityaugmented airframe is:

Referring to FIG. 10, in accordance with one embodiment of theinvention, the rate gyro 20 provides a first input error signal forcontrol of a voltage applied to a flapper valve core 112 of the actuator16. A second error signal is applied by the pilot actuating the control10 which moves the tap 114 on a potentiometer 115 to produce an errorvoltage at terminal 114b. A third source of error voltage is from thefeedback transducer 50 connected to terminal 116. A fourth control 1 17is provided for balance purposes.

The gyro 20 is provided with an output transformer primary which iscoupled to a secondary 121 connected across terminals 122 and 123.Terminals 122 and 123 are connected together at terminal 124 by way ofresistors 125 and 126. The terminal 114 b is connected to the junctureof secondary 121 and resistor 125. The signal at terminal 123 is appliedby way of a capacitor to the input terminal 131 of an amplifier 132 A DCsignal representative of the voltage at tap 114 and appearing acrosscapacitor 99 is applied by way of resistors 125 and 126 to the centertap 135 of an input transformer 137 of a bilateral gate 136. The primarywinding of the input transformer 137 is shunted by a diode 128. Oneterminal of the primary winding is connected by way of resistor 139 toterminal 140, the plus terminal of a DC supply source (not shown). Theouter terminal of the primary winding of the transformer 137 isconnected to the cathode electrode of a transistor 178 of a pulsegenerator 98. The pulse generator 98 is excited by a 1 l5 volt, 400cycle per second supply; as a result, the generator 98 produces negativegoing pulses at the rate of 400 pulses per second. The gate pulse isphased so that it opens gate 136 during the peak of the negativehalf-cycle for a period of 0.002 seconds. During this interval, a pulsecurrent circulates in a loop which includes Zener diode 145, diode 146,diode 147 and Zener diode 148. Dlodes 146 and 147 are connected atopposite polarities as are diodes and 148. The common juncture 150 inthe loop is connected by way of a capacitor 151 to ground and to thebase electrode of transistor 152. The collector electrode of transistor152 is connected to the terminal 140 by way of a resistor 141. Thecollector voltage is maintained constant by means of a Zener diode 110.The emitter is connected by way of resistor 153 to ground and by way ofseries capacitor 154 and resistor 155 to the input terminal 131 of theamplifier 132.

By this means, the voltage change produced by actuation of the tap 114by the pilot adjusting the control element 10 is sampled and stored oncapacitor 151. The combination of resistors 125 and 126 develops a timeconstant for the desired lag in the response of the system to changesintroduced by the pilot at potentiometer tap 1 14.

The gate 136 serves to either charge or discharge capacitor 151 at arate controlled by the magnitude and direction of change in the positionof tap 114. The circuit comprising capacitor 154 and resistor 155transmits the unbalance signal from the sample and hold circuitinvolving capacitor 151 to the input of amplifier 132; it transmits onlythe AC components of the unbalanced signal. Capacitor 154 and resistor155 comprise the washout circuit described previously. When the pilotshifts the control stick to a difierent position, the voltage applied toterminal 131 is a measure of the magnitude and direction of the changein position. The time constant of the circuit comprising capacitor 154and resistor 155 is of the order of 2 to 3 seconds.

The signal from the feedback transducer 50 is applied by way of anattenuator pad to terminal 131. A balance voltage is applied from apotentiometer 117 and also by way of an attenuating pad 162 and aresistor 163 to terminal 131. Terminal 131 is connected to ground by wayof a resistor 164. With no unbalance signal present in the system, thevoltage from potentiometer 117 is applied to terminal 131 in an amountexactly equal to the DC bias applied to an input terminal 165 ofamplifier 132.

In the pulse generator 98, the 400 cycle voltage is applied by way ofresistor 170 to a Zener diode 171 which clips the peaks of the voltagewave and maintains a relatively constant reference voltage. The positivegoing waves are passed by way of diode 172 to the base electrode of atransistor 173. The transistor 173 generates a spacer pulse which isapplied to transistor 174 by way of diode 175. The time delay is suchthat a positive going pulse of 0.0002 seconds is generated at thecollector of transistor 174 and applied by way of a line 176 and diode177 to the input terminal 131 of the amplifier 132 of fixed high gain.The positive going pulse is also applied to the base of a transistor 178whose output is a negative going pulse of 0.0002 seconds duration. Thelatter pulse is applied by way of a line 179 from the cathode electrodeof the transistor 178 and by way of diode 180 to the input terminal 165.

The-positive and negative going pulses applied to amplifier 132 serve toactuate the amplifier 132 for relatively short intervals during eachcycle. For example, the period of the 400 cycle input signal is 2.5milliseconds whereas the gating pulses centered at the positive andnegative half cycles are 0.0002 seconds. The unbalance signal which maybe applied to terminal 131 from either the gyro 20, the potentiometer114, or the feedback transducer 50 will thus be amplified in amplifier132. Amplifier 132 comprises three stages 191, 192 and 193. The emitterelectrode of the output stage 193 is connected by way of channel 194 tothe center tap of an input transformer 195 of a bilateral gate 196. Gate196 is in all essential characteristics the same as gate 136 including asample and hold capacitor 197. The voltage across the capacitor 197 isapplied to the input of a Darlington connected pair of transistors 199and 200. The collector electrode of the output transistor 200 isconnected by way of resistor 201 to one terminal of the torque motorwinding 112. The other terminal of the torque motor winding 112 isconnected by way of a resistor 202 to the emitter electrode oftransistor 193 and by way of resistor 203 to ground. Negative gatingpulses are applied to the gate 196 by way of conductor 204. The circuitoperates such that during the period of the negative going pulses onchannel 204 a voltage is stored on capacitor 197. This voltage is thencompared with the voltage appearing at the emitter of transistor 193during the remainder of each cycle The difference in voltage isrepresentative of the current which flows through the torque motorwinding 112. This voltage may be either positive or negative and thusmay be employed to control a flapper in a conventional hydraulicactuator. When no current flows through the torque motor winding 112,the flapper in the first stage of a hydraulic actuator, such as actuator16, is in a balanced position. Positive or negative current flowing inthe winding 112 will change the position .of the flapper in thehydraulic actuator to cause the actuator to make a suitable change inthe control system.

It will now be seen that the compensating networks 24 and 38 of FIG. 1have been consolidated in FIG. 10. That is, the command signal from tap114 and the error signal from the gyro 20 pass through the same lagnetwork wherein the lag is controlled by condenser 151. They passthrough a first gating amplifier stage 152 and thence through thewashout circuit which includes condenser 154 and resistor 155. Terminal131 represents the summation point for signals from the washout circuitand from the feedback actuator transducer 16, the signal from the lattertransducer being applied to terminal 1 16.

The use of one lag circuit and one washout circuit for both the commandsignal and the error signal assures identical phase and amplituderesponses and permits introduction of a LII command from lever 10without compromise or alteration of the gain of amplifier 132.

Where desired, separate lag and washout circuits may be employed.

Further, only one channel has been shown, it being understood that sucha channel will be provided for each of the yaw, roll and pitch axes in acomplete stability'control augmentation system.

While a preferred embodiment of the invention, together withmodifications thereof, has been described in detail herein and shown inthe accompanying drawing, it will be evident that various furthermodifications are possible in the arrangement, and construction of itscomponents without departing from the scope of the invention.

lclaim:

1. An aircraft control system for selectively establishing andmaintaining an attitude rate in the presence of disturbing forces fromwind gusts, weapon recoil, and the like, which comprises:

a. a control loop including said aircraft, a reference element coupledto said aircraft, first signal coupling means connected to saidreference element, a fixed high gain amplifier for amplifying errorsignals from said coupling means and an actuator responsive to signalsfrom said amplifier for continuous control of said attitude rate;

b. a manual input element for introducing a command for a desired changein attitude rate including transducer means to produce a command signal,and

c. second coupling mans for applying said command signal with said errorsignal for amplification in said amplifier to produce a specificaircraft attitude rate response independent of the response of saidcontrol loop to said disturbing forces.

2. The combination set forth in claim 1 wherein said coupling meansapply said command signal and said error signal to the input to saidamplifier with time delays and frequency responses which bearpredetermined relationships to each other.

3. The combination set forth in claim 2 wherein said coupling means havethe same time delays and the same frequency responses.

4. The combination set forth in claim 2 in which said coupling means arecommon to the signal paths for both said error signal and said commandsignal.

5. In a control and stabilization system for a vehicle, the combinationcomprising:

a feedback system having an outer feedback loop which includes saidvehicle and responsive to external forces applied to thereto to providestabilization relative to an established course, said feedback systemalso including an inner control feedback loop responsive to the amountof stabilizaiton provided by said first feedback loop, and

a feedforward control loop including a command element and meansresponsive to commands applied to the vehicle and having a fixed highgain amplifier means common to said inner loop for controlling thecourse of travel of the vehicle, and

a compensating network having a lag and washout circuit in said outerfeedback loop which also forms a lag and washout circuit in saidfeedforward control loop.

6. A control and stabilization system for a vehicle, the combinationcomprising:

pilot control means for controlling the direction of travel of saidvehicle,

actuating means having a mechanical input proportional to a desired rateof change in direction of travel of said vehicle coupled to said controlmeans for establishing the position of said actuating means,

first transducer means for detecting the rateof divergence of saidvehicle from an established direction of travel,

a fixed high gain amplifier completing a feedback loop between saidtransducer means and said actuating means for applying a signal to saidactuating means proportional to rate of deviation of said vehicle fromits established direction of travel to restore said vehicle to saidestablished direction of travel,

means responsive to the output of said actuating means and coupled tosaid amplifier for returning the output of said amplifier to zero, and

second transducer means responsive to said control means and connectedto said amplifier for applying to said amplifier a signal proportionalto a desired rate of change in direction of said vehicle.

7. A control and stabilization system as set forth in claim 6 whereinsaid detecting means includes a rate gyro coupled to said vehicle andwherein a compensating network is respon' sive to the output of saidfirst transducer and to said second transducer and tailors the transferfunction of said feedback loop to be greater than the transfer functionof said vehicle throughout the significant control/response range ofsaid vehicle.

8. A system for control and stabilization of a helicopter having controlsurfaces, the combination comprising:

a pilot actuated lever for generating a signal proportional to a desiredrate of change in direction of travel,

an actuator responsive to movement of said lever for generating amechanical movement proportional to the signal generated by said lever,

a control system coupled to the control surfaces of said helicopter andresponsive to the output of said actuator,

a rate gyro coupled to said helicopter for generating a signalproportional to the rate of deviation of said helicopter from areference direction of travel,

a feedback transducer responsive to the output of said actuator forgenerating a signal proportional to the output thereof,

a command signal transducer coupled to said lever for generating asignal proportional to the movement thereof, and

fixed high gain amplifier means coupled at its input to the outputs ofsaid rate gyro, said feedback transducer, and said command transducerfor applying a signal to said actuator proportional to the sum of thesignals applied to said input.

9. The method of control of attitude rate of a craft which comprises:

a. generating a high constant gain feedback signal dependent uponoutside disturbances on said craft,

b. generating a control signal by manual actuation of a control element,

0. generating a feedforward signal in response to said control signal,and

d. mixing said feedforward signal and said feedback signal with liketime lag and frequency characteristics to modify the change rate withoutalteration of the response of said feedback loop to outsidedisturbances.

10. A control and stabilization system for an aircraft which comprises:

a. an airframe,

b. pilot steering means,

c. a stability feedback loop including means for generating a firstsignal dependent upon the rate of movement of the aircraft relative to agiven axis, control surface actuating means connected to be mechanicallyresponsive to said steering means and fixed high gain means for treatingsaid first signal for driving said actuating means in accordance with atransfer function greater than the transfer function of said airframethrough the significant control/response range of said airframe, and

d. means for generating a second signal in response to movement ofsaidsteering means, and

e. coupling means for applying said second signal to said fixed highgain means in said stability loop to produce an airframe rate responseabout said given axis proportionate to movement of said pilot controlinput means throughout the significant control/response range of saidairframe.

11. In an aircraft stability and control augmentation system, thecombination comprising:

a. an airframe of controllable attitude,

b. a control system including an attitude stabilizing feedback loopincluding airframe attitude rate detecting means, a first signalgenerating means responsive to said attitude rate detecting means, andactuating means operative to control attitude of said airframe andresponsive to said first signal generating means,

c. fixed high gain amplifier means in said stabilizing feedback loop toestablish gain in said loop throughout the significant control/responserange of said airframe sufficient to dominate the responsecharacteristics of the airframe to outside disturbances within saidrange,

d. a pilot input means including pilot steering means, and a secondsignal generating means responsive to changes in position of saidsteering means,

e. means connecting said second signal generating means to saidamplifier means,

f. means connecting said actuating means to said steering means forproducing a proportionate airframe rate response throughout thesignificant control/response range of said airframe to movement of saidsteering means.

12. A system for control and stabilization of a helicopter havingcontrol surfaces, the combination comprising:

a pilot actuated lever for generating a signal proportional to a desiredrate of change in direction of travel,

an actuator response to movement of said lever for generating amechanical movement proportional to the signal generated by said lever,

a control system coupled to the control surfaces of said helicopter andresponsive to the output of said actuator,

a rate gyro coupled to said helicopter for generating a signalproportional to the rate of deviation of said helicopter from areference direction of travel,

a feedback transducer responsive to the output of said actuator forgenerating a signal proportional to the output thereof,

a command signal transducer coupled to said lever for generating asignal proportional to the movement thereof,

a time lag circuit coupled at its input to the output of said gyro andsaid command transducer,

first amplifier means connected to said lag circuit,

a washout circuit connected to the output of said first amplifier means,and

second amplifier means connected at its input to the output of saidwashout circuit and to the output of said feedback transducer forapplying a feedback signal to said actuator proportional to the sum ofthe signals applied to said input.

13. A helicopter control and stabilization system as set forth in claim12 including:

a first bilateral gate connected to the input of said first amplifier,

a second bilateral gate connected to the output of said secondamplifier, and

a synchronizing pulse generator connected to said first bilateral gate,to the input of said second amplifier, and to said second bilateralgate.

14. A helicopter control and stabilization system as set forth in claim13 wherein said pulse generator includes means for generating negativegoing pulses connected to said first and second bilateral gates and theinput of said second amplifier, and means for generating positive goingpulses connected to the input of said second amplifier.

15. A helicopter control and stabilization system as set forth in claim13 wherein said lag circuit includes two parallel resistors connected toone side of said first bilateral gate and a capacitor connected to theopposite side of said first bilateral .8??? 16. A helicopter control andstabilization system as set forth Page 1 of 2 UNITED STATES PATENTOFFICE CERTIFICATE OF CORRECTION Patent No. 3,58 4, 81 Dated June 15,1971 ln'ventor(s) Milford R. Murphy It is certified that error appearsin the above-identified patent and that said Letters Patent are herebycorrected as shown below:

Col. 1, line 63, change "disturbance" to --disturbances.

Col. 2, line 30, change "above system" to --above control system-.

Col. 3, line 6 change "determined 'the design to --determined by thedesign-;

line 67 change the formula so that it reads as follows:

line 68 change the formula so that it reads as follows:

line 72 change "functions" to --function; line 7" change "determined thebasic" to --determined by the basic-. Col. 4, line 18 change "the" to--then--;

line 68 I change "tow" to two-; line 75 change "consideration" to--considerations-.

Col. 5, line 32 change "13" to -l/H-.

Col. 6, line 28 after "network 2 4" and before "in" delete "is";

line 4% change "portions" to -portion--;

line change- "network 66 is the" to -network M5 in the. J

Page 2 of 2 P0405) UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTIONPatent No. Dated n 15, 1971 Inventor(s) Milford R. Murphy It iscertified that error appears in the above-identified patent and thatsaid Letters Patent are hereby corrected as shown below:

Col. 7, line 17, After "G change "function" to --invo1ves-',

line 7", change "transfer functions" to v transfer function.

Col. 8, line 1 change the formula so that it reads as follows: a

G G -0.6M(12.5s +1) line 28 after "amplifier 132" insert line 36 change"outer" to -other--;

, line 67 change "AC components" to -AC component.

Col. 9, line 51 after "cycle" insert line 58 change "current flowing" to--currents flowing--.

Col. 10 line 18 after "attitude rate" insert reference-;

' line 30 change "coupling mans" to --coupling means;

line 51 after "applied and before "thereto" delete "to" Signed andsealed this 9th day of Noirember 1971.

(SEAL) Attest:

EDWARD M.FLETCHER,JR. Attesting Officer ROBERT GOTTSGHALK J ActingCommissioner of Patents

1. An aircraft control system for selectively establishing andmaintaining an attitude rate in the presence of disturbing forces fromwind gusts, weapon recoil, and the like, which comprises: a. a controlloop including said aircraft, a reference element coupled to saidaircraft, first signal coupling means connected to said referenceelement, a fixed high gain amplifier for amplifying error signals fromsaid coupling means and an actuator responsive to signals from saidamplifier for continuous control of said attitude rate; b. a manualinput element for introducing a command for a desired change in attituderate including transducer means to produce a command signal, and c.second coupling mans for applying said command signal with said errorsignal for amplification in said amplifier to produce a specificaircraft attitude rate response independent of the response of saidcontrol loop to said disturbing forces.
 2. The combination set forth inclaim 1 wherein said coupling means apply said command signal and saiderror signal to the input to said amplifier with time delays andfrequency responses which bear predetermined relationships to eachother.
 3. The combination set forth in claim 2 wherein said couplingmeans have the same time delays and the same frequency responses.
 4. Thecombination set forth in claim 2 in which said coupling means are commonto the signal paths for both said error signal and said command signal.5. In a control and stabilization system for a vehicle, the combinationcomprising: a feedback system having an outer feedback loop whichincludes said vehicle and responsive to external forces applied tothereto to provide stabilization relative to an established course, saidfeedback system also including an inner control feedback loop responsiveto the amount of stabilizaiton provided by said first feedback loop, anda feedforward control loop including a command element and meansresponsive to commands applied to the vehicle and having a fixed highgain amplifier means common to said inner loop for controlling thecourse of travel of the vehicle, and a compensating network having a lagand washout circuit in said outer feedback loop which also forms a lagand washout circuit in said feedforward control loop.
 6. A control andstabilization system for a vehicle, the combination comprising: pilotcontrol means for controlling the direction of travel of said vehicle,actuating means having a mechanical input proportional to a desired rateof change in direction of travel of said vehicle coupled to said controlmeans for establishing the position of said actuating means, firsttransducer means for detecting the rate of divergence of said vehiclefrom an established direction of travel, a fixed high gain amplifiercompleting a feedback loop between said transducer means and saidactuating means for applying a signal to said actuating meansproportional to rate of deviation of said vehicle from its establisheddirection of travel to restore said vehicle to said establisheddirection of travel, means responsive to the output of said actuatingmeans and coupled to said amplifier for returning the output of saidamplifier to zero, and second transducer means responsive to saidcontrol means and connected to said amplifier for applying to saidamplifier a signal proportional to a desired rate of change in directionof said vehicle.
 7. A control and stabilization system as set forth inclaim 6 wherein said detecting means includes a rate gyro coupled tosaid vehicle and wherein a compensating network is responsive to theoutput of said first transducer and to said second transducer andtailors the transfer function of said feedback loop to be greater thanthe transfer function of said vehicle throughout the significantcontrol/response range of said vehicle.
 8. A system for control andstabilization of a helicopter having control surfaces, the combinationcomprising: a pilot actuated lever for generating a signal proportionalto a desired rate of change in direction of travel, an actuatorresponsive to movement of said lever for generating a mechanicalmovement proportional to the signal generated by said lever, a controlsystem coupled to the control surfaces of said helicopter and responsiveto the output of said actuator, a rate gyro coupled to said helicopterfor generating a signal proportional to the rate of deviation of saidhelicopter from a reference direction of travel, a feedback transducerresponsive to the output of said actuator for generating a signalproportional to the output thereof, a command signal transducer coupledto said lever for generating a signal proportional to the movementthereof, and fixed high gain amplifier means coupled at its input to theoutputs of said rate gyro, said feedback transducer, and said commandtransducer for applying a signal to said actuator proportional to thesum of the signals applied to said input.
 9. The method of control ofattitude rate of a craft which comprises: a. generating a high constantgain feedback signal dependent upon outside disturbances on said craft,b. generating a control signal by manual actuation of a control element,c. generating a feedforward signal in response to said control signal,and d. mixing said feedforward signal and said feedback signal with liketime lag and frequency characteristics to modify the change rate withoutalteration of the response of said feedback loop to outsidedisturbances.
 10. A control and stabilization system for an aircraftwhich comprises: a. an airframe, b. pilot steering means, c. a stabilityfeedback loop including means for generating a first signal dependentupon the rate of movement of the aircraft relative to a given axis,control surface actuating means connected to be mechanically responsiveto said steering means and fixed high gain means for treating said firstsignal for driving said actuating means in accordance with a transferfunction greater than the transfer function of said airframe through thesignificant control/response range of said airframe, and d. means forgenerating a second signal in response to movement of said steeringmeans, and e. coupling means for applying said second signal to saidfixed high gain means in said stability loop to produce an airframe rateresponse about said given axis proportionate to movement of said pilotcontrol input means throughout the significant control/response range ofsaid airframe.
 11. In an aircraft stability and control augmentationsystem, the combination comprising: a. an airframe of controllableattitude, b. a control system including an attitude stabilizing feedbackloop including airframe attitude rate detecting means, a first signalgenerating means responsive to said attitude rate detecting means, andactuating means operative to control attitude of said airframe andresponsive to said first signal generating means, c. fixed high gainamplifier means in said stabilizing feedback loop to establish gain insaid loop throughout the significant control/response range of saidairframe sufficient to dominate the response characteristics of theairframe to outside disturbances within said range, d. a pilot inputmeans including pilot steering means, and a second signal generatingmeans responsive to changes in position of said steering means, e. meansconnecting said second signal generating means to said amplifier means,f. means connecting said actuating means to said steering means forproducing a proportionate airframe rate response throughout thesignificant control/response range of said airframe to movement of saidsteering means.
 12. A system for control and stabilization of ahelicopter having control surfaces, the combination comprising: a pilotactuated lever for generating a signal proportional to a desired rate ofchange in direction of travel, an actuator response to movement of saidlever for generating a mechanical movement proportional to the signalgenerated by said lever, a control system coupled to the controlsurfaces of said helicopter and responsive to the output of saidactuator, a rate gyro coupled to said helicopter for generating a signalproportional to the rate of deviation of said helicopter from areference direction of travel, a feedback transducer responsive to theoutput of said actuator for generating a signal proportional to theoutput thereof, a command signal transducer coupled to said lever forgenerating a signal proportional to the movement thereof, a time lagcircuit coupled at its input to the output of said gyro and Said commandtransducer, first amplifier means connected to said lag circuit, awashout circuit connected to the output of said first amplifier means,and second amplifier means connected at its input to the output of saidwashout circuit and to the output of said feedback transducer forapplying a feedback signal to said actuator proportional to the sum ofthe signals applied to said input.
 13. A helicopter control andstabilization system as set forth in claim 12 including: a firstbilateral gate connected to the input of said first amplifier, a secondbilateral gate connected to the output of said second amplifier, and asynchronizing pulse generator connected to said first bilateral gate, tothe input of said second amplifier, and to said second bilateral gate.14. A helicopter control and stabilization system as set forth in claim13 wherein said pulse generator includes means for generating negativegoing pulses connected to said first and second bilateral gates and theinput of said second amplifier, and means for generating positive goingpulses connected to the input of said second amplifier.
 15. A helicoptercontrol and stabilization system as set forth in claim 13 wherein saidlag circuit includes two parallel resistors connected to one side ofsaid first bilateral gate and a capacitor connected to the opposite sideof said first bilateral gate.
 16. A helicopter control and stabilizationsystem as set forth in claim 15 wherein said washout circuit includes acapacitor in series with a resistor connected to the output of saidfirst amplifier and the input of said second amplifier.